Methods of controlling thrust in a rocket motor

ABSTRACT

A propulsion thrust control system and method for controlling thrust in a rocket motor includes configuring valves of an energized rocket motor to an initial total valve area according to a total thrust command. The total thrust command is converted into a commanded propellant mass flow discharge rate. A varying total valve area is computed from an error between the commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate. The valves are reconfigured according to a distribution of the varying total valve area. The propulsion system includes a pressure vessel with valves and a controller for regulating the valve area according to a propellant mass flow discharge rate from the pressure vessel.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a divisional of U.S. patent application Ser. No.11/366,252, filed Mar. 2, 2006, pending, the entire disclosure of whichis hereby incorporated herein by this reference.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates to a propulsion thrust control systemconfigured for control of a rocket-propelled vehicle.

2. State of the Art

Solid propellant rocket motors employ a propellant comprising a solidfuel charge or “grain” which burns to generate exhaust gases and othercombustion products, which are expelled through one or more nozzles ofthe rocket motor to provide thrust. Once a grain of solid propellant isignited, it is difficult to extinguish and the entire grain isordinarily consumed after ignition. Additionally, effecting variation ofthrust is more difficult in solid propellant than in liquid propellantrocket engines. However, simple structural design of solid propellantrocket motors and ease of storage of the solid propellant are advantagesof the solid propellant motor.

Attitude control, in the form of influencing the pitch, yaw, and/or rollof the rocket assembly in flight, may be accomplished with a thrustvector control (TVC) system or a separate attitude control system (ACS).A TVC system may comprise an axial thrust nozzle rotationallypositionable at a desired angle within a range offset from thelongitudinal axis of the rocket motor to alter the vector at which thecombustion products exit the rocket motor. Repositioning of the nozzlealters the direction of the forces acting on the vehicle in which therocket motor is installed to alter the vehicle's direction of flight.Single, movable TVC nozzles provide adequate control over the rocketassembly's yaw and pitch, but do not provide any significant degree ofroll control.

Multiple rocket engines or gas generators and associated thrusters areoften employed to control attitude. The rocket engines or thrusters areoffset from the longitudinal axis of the rocket motor assembly so thatfiring of selected ones or groups of the engines or powering of selectedones or groups of thrusters enables attitude control over the rocketmotor assembly. Use of a separate ACS in combination with one or moreaxial thrust engines or thrusters increases the weight of the rocketmotor assembly due to the additional hardware. A separate ACS may use asolid-propellant gas generator directly connected to a manifoldproviding a selective hot gas flow to nozzle valve clusters. Rollcontrol may be provided by the ACS or through the inclusion of aseparate roll control system (RCS). Separate gas generators andthrusters may be provided for the RCS.

As stated, each of the control systems such as the ACS and the TVCdirect gases through valves that results in the generation of thrust foraltering the vehicle's direction of flight. Valves are known that may beconfigured in either “open” or “closed” states. Additionally, valvesthat include at least a third or “partially open” state are known as“proportional valves.” The state of each valve is determined by thecross-sectional area of the orifice or “throat” for each valve. Thestate of each of the valves is controlled by a motor controller orcontrol system, which adjusts the state of the various valves based uponone or more control inputs.

One method for controlling the state of each of the valves relies uponpressure measurements which become the control inputs or processvariables to the control system. The control inputs in this mode ofcontrol are net thrust per valve set (F_(net,i)) and a pressure setpoint or limit. As used herein, net thrust is defined as thedifferential thrust between opposing valve pairs.

According to a thrust/pressure-only control methodology, the totalthroat area is regulated to obtain a pressure. An approximate throatarea (control variable) for a pressure command is given in the followingequation:

A _(t)=(C*· A _(s) r(P _(cmd) /P _(ref))^(n))/(g _(c) P _(cmd))   [1]

where: ρ=density of propellant [lbm/in³]

-   -   P_(ref)=reference pressure used to determine r [lbf/in²]    -   A_(s)=surface area of propellant [in²]    -   n=exponent    -   r=burn rate @ P_(ref) [in/sec]    -   A_(t,i)=throat flow area of valve i [in²]    -   P_(cmd)=gas generator pressure command [lbf/in²]    -   g_(c)=gravitational constant [(lbm/lbf)(ft/s²)]    -   C*=characteristic exhaust velocity [ft/sec]        where the pressure command could be generated by the following        equation:

P _(cmd)=max {C ₁ max(F _(net,i)), C ₂ max(F _(net,i))^(1/n)}  [2]

where: C₁=1/(A_(t(FULL OPEN)) C_(F))

-   -   C₂=P_(ref) {1/(I_(SP) ρ A_(S) r)}^(1/n)

where: I_(SP)=specific impulse [s]

-   -   C_(F)=discharge coefficient of valve

The pressure command is determined by taking the maximum of twocalculations. The first involves the pressure required to meet the worstcase (highest) net thrust command relative to the available throat areafor a given valve set. This pressure is determined by the specific valvecharacteristics. The second involves the pressure required to meet allnet thrust requirements. This pressure is determined by propellantcharacteristics and geometry,

The net thrust from a given valve set is proportional to the gasgenerator (GG) pressure and difference in regulated throat areas ofopposing valves:

F _(net,i)=(A _(t,i) −A _(t,j))P C _(F), where valves i and j comprise avalve set and provide thrust in opposite directions.   [3]

The distribution of the regulated throat area meets thrust commands inaccordance with the system requirements. By way of simplified exampleand to avoid infinite solutions of the pressure-only controlmethodology, the throat areas are evenly distributed (in proportion toflow capacities) among all valves and the same offset, in oppositedirections, is added to each valve (in a valve set) to achieve the netthrust command. It is known that the faster response times of the valvescompared to slower response of the gas generator, allow controlling netthrust during transient events in the gas generator. This is especiallytrue, for example, when controlling ACS valves that are much smallerthan divert valves since ACS valves have a much smaller effect on gasgenerator (GG) pressure.

While the thrust control approach using a pressure-only methodology maybe effective, inefficiencies remain, specifically, the sensitivity tothe GG variation. For example, if the propellant is not burning at thespecifically designed burn rate or propellant surface area deviates fromthat designed, a differential (i.e., more or less) mass flow results.For example, assume that more mass is generated than desired because theactual burn rate is higher than designed. If pressure is regulated, thevalves will have to open more than predicted to accommodate the largeractual mass flow generated per unit time (“m-dot_(gen)”), resulting inan excess (i.e., wasted) thrust (i.e., more mass flow discharged perunit time (“m-dot_(disch)”) than required to meet all the net thrustcommands) as shown by:

Σm-dot_(disch) =P ΣA _(t, i) g _(c) /C*>(ΣF _(net, 1))/Isp=m-dot_(cmd).  [4]

As an example, demand for thrust in a specific direction occurs bycommanding or controlling a valve pair to achieve the commanded thrust.However, the aggregate throat area of all the valves must compensate forpressure regulation, while each opposing pair has the correctdifferential throat area. Inherently, the pressure compensation requiresopening more than predicted resulting in wasted propellant by dumpingsome of the generated gas. Propulsion systems that incorporate solidpropellant, while desirable due to performance and weight reduction,become less desirable as inefficiencies are introduced due to dumping ofgases that could otherwise be used for motion control.

For thrust control using a pressure-only methodology, an updated burnrate coefficient (r_(update)) can be calculated whenever the system isin a quasi-steady-state condition, as defined when the pressure iswithin a defined percentage of the commanded value and oscillates belowa threshold value. In this condition, an updated burn rate coefficientmay be simply calculated by knowing measured gas generator (GG) pressure(P_(measured)) and measured total throat areas (A_(t, measured, i))achieving that pressure with all system parameters assumed constant asshown by:

r _(update) =ΣA _(t,measured, i) g _(c) P _(measured)/(C*ρ A _(s)(P_(measured) /P _(ref))^(n)).   [5]

Knowledge of the actual burn rate coefficient helps to predictpropellant burn-out, but does not directly aid in thrust control whenimplementing a thrust control methodology based on pressure-only controlinputs. Additionally, while a pressure command change (ΔP(=P_(nom)−P_(update)) can be subtracted from the pressure command:

P _(nom)=(A _(s) ρ r _(nom) C*/(ΣA _(t,nom) P ^(n) _(ref) g_(c)))^((1/(n−1))   [6]

P _(update)=(A _(s) ρ r _(update) C*(ΣA _(t,nom) P ^(n) _(ref) g_(c)))^((1(n−1)))   [7]

P _(cmd) =P _(cmd)(P _(nom) −P _(update))   [8]

to compensate for the burn rate variability with P_(nom) representingthe pressure at a nominal burn rate coefficient and nominal total throatarea and P_(update) representing the pressure at the updated burn rateand nominal total throat area, an updated ΔP must be calculated any timea new estimate of the burn rate coefficient occurs or any time thepressure command changes.

In view of the above-enumerated deficiencies in the state of the artwith respect to pressure-only thrust control of a rocket-propelledvehicle, it would be desirable to develop a methodology for controllingthrust in a vehicle for improving the inefficiencies and for calculatingthrust control commands.

BRIEF SUMMARY OF THE INVENTION

A propulsion thrust control system and method for controlling thrust ina rocket motor is provided. In one embodiment of the present invention,a method of controlling thrust in a rocket motor includes configuringvalves of an energized rocket motor to an initial total valve areaaccording to a total thrust command. The total thrust command isconverted into a commanded propellant mass flow discharge rate. Avarying total valve area is computed from an error between the commandedpropellant mass flow discharge rate and a calculated or measuredpropellant mass flow discharge rate. The valves are configured accordingto a distribution of the varying total valve area and net thrustcommands at every computational cycle.

In another embodiment of the present invention that minimizes wastedpropellant (i.e., optimized efficiency), a method for controlling thrustin a rocket motor is provided. The method includes receiving net thrustcommands for orienting a rocket motor propelled vehicle and calculatingfrom the net thrust commands a minimum total thrust required from a gasgenerator of the rocket motor. The total valve area is continuouslyreconfigured according to the total thrust and in response to acomparison of a commanded propellant mass flow discharge rate and acalculated propellant mass flow discharge rate, as shown byΣm-dot_(disch)=F_(tot.min)/I_(sp).

In a further embodiment of the present invention, a propulsion system isprovided. The propulsion system includes a pressure vessel containing apropellant and at least one axial thrust valve in communication with thepressure vessel and configured for selectively releasing gases generatedby combustion of the propellant within the pressure vessel to provideaxial thrust. The propulsion system further includes at least onemaneuver control valve in communication with the pressure vessel andconfigured for selectively releasing gases generated by combustion ofthe propellant within the pressure vessel to provide thrust formaneuvering and a controller for regulating a valve area of the at leastone axial valve and at least one maneuver control valve according to apropellant mass flow discharge rate from the pressure vessel.

BRIEF DESCRIPTION OF THE SEVERAL VIEWS OF THE DRAWINGS

The foregoing and other advantages of the invention will become apparentupon reading the following detailed description and upon reference tothe drawings in which:

FIG. 1 shows an exemplary rocket motor embodiment in longitudinalcross-section, in accordance with an embodiment of the presentinvention;

FIG. 2 is a schematic axial view of maneuver control valves and maneuvercontrol thrusters of a rocket motor of FIG. 1, in accordance with anembodiment of the present invention;

FIG. 3 is a schematic diagram of a rocket motor, in accordance with anembodiment of the present invention;

FIGS. 4A and 4B are flowcharts illustrating calculation of total thrustcommands, in accordance with an embodiment of the present invention;

FIG. 5 is a flowchart for initially configuring control valves, inaccordance with an embodiment of the present invention;

FIG. 6 is a flowchart illustrating two exemplary modes for configuringcontrol valves into a steady state configuration, in accordance with oneor more embodiments of the present invention;

FIG. 7 is a flowchart illustrating automatic control of control valvesaccording to a mass flow methodology, in accordance with an embodimentof the present invention; and

FIG. 8 is a flowchart illustrating adjustment of a burn rate, inaccordance with an embodiment of the present invention.

DETAILED DESCRIPTION OF THE INVENTION

One exemplary embodiment of a rocket motor according to the presentinvention, which may comprise an upper or final stage rocket motor, isdepicted in FIG. 1. The motor case assembly comprises a motor casehousing 12 which houses the pressure vessel 14 (also sometimes termed a“motor case”) having a plurality of valves in communication therewith.Within the pressure vessel 14, low density foam 20 surrounds andinsulates the solid propellant grain 22. In one exemplary, nonlimitingimplementation of the present invention, the motor case assembly withinmotor case housing 12 may have a diameter 16 of between about 25 and 30inches, currently preferred to be 27.6 inches and a length 18 of between30 and 35 inches, currently preferred to be 32 inches.

Solid propellant grain 22 may comprise, for example, a free standingclass 7 HMX (cyclotretramethylenetetranitramine)-oxidized compositepropellant with a binder system based on hydroxyl-terminatedpolybutadiene (HTPB) polymer and cured with isophorene diisocyanate(IPDI) curative including a small amount of carbon black as anopacifier, the propellant being formulated to burn stably over a widepressure range. Alternatively, solid propellant grain 22 may comprise,for example, an aluminum powder-fueled, hydroxyl-terminatedpolybutadiene (HTPB) polymer-based binder. One currently preferredpropellant is a nonaluminized HTPB propellant grain of 228 lbm for theabove-sized rocket motor. The solid propellant chosen for use may be anyof those known to one of ordinary skill in the art, as the presentinvention does not require a specific propellant for implementation.

The axial thrust valve 10 may comprise a pintle valve configured forproportional operation and control of axial thrust through axialthruster 26, which may be configured, by way of example only, to providea maximum of 4,000 lbf of thrust. As best observable from FIG. 2, whichillustrates the exit cone 24 of axial thruster 26 in broken lines forclarity, maneuver control thrusters 32, 34, 40 a, 40 b, 42 a, 42 b arerespectively operably coupled to maneuver control valves 28, 30, 36 a,36 b, 38 a, 38 b and located and oriented to effect maneuveringfunctions including pitch, yaw and roll control. Maneuver control valves28, 30, 36 a, 36 b, 38 a, 38 b may comprise proportional valves.

As depicted in FIG. 2, selective operation of two maneuver controlvalves 28, 30 with respectively associated coplanar maneuver controlthrusters 32, 34 located 180° apart and oriented transverse to thelongitudinal axis L of the rocket motor may be used for pitch control.Yaw control may be effected by selective operation of either pairedmaneuver control thrusters 40 a and 42 a by maneuver control valves 36 aand 38 a or paired and diametrically opposed maneuver control thrusters40 b and 42 b by maneuver control valves 36 b and 38 b. As shown, pairedmaneuver control thrusters 40 a and 42 a and 40 b and 42 b are coplanar,oriented transverse to longitudinal axis L of the rocket motor and maybe used to provide balanced, parallel thrust vectors to either side oflongitudinal axis L at identical lateral offsets therethrough.

Roll control may be effected by selectively using two sets of maneuvercontrol valves 36 a, 36 b, 38 a, 38 b and respectively associatedcoplanar maneuver control thrusters 40 a, 40 b, 42 a, 42 b. Roll in afirst rotational direction may be effected by opening maneuver controlvalves 36 a and 36 b to power maneuver control thrusters 40 a and 40 band cause them to provide a first set of opposing but complementarythrust vectors laterally offset from longitudinal axis L, while roll ina second direction may be effected by opening maneuver control valves 38a and 38 b to power maneuver control thrusters 42 a and 42 b and causethem to provide a second set of opposing but complementary thrustvectors laterally offset from longitudinal axis L.

Increasing the total flow area by opening any of the aforementionedvalves during combustion of solid propellant grain 22 will necessarilydecrease pressure within the pressure vessel 14. This will reduce theburn rate of the propellant and, therefore, diminish thrust. In anexemplary embodiment of the invention, when the axial thrust valve 10 isin the fully open position, and all other valves are closed, theinternal pressure within pressure vessel 14 may be such that the solidpropellant grain 22 will have the lowest possible steady state burnrate, which corresponds to the minimum desired thrust. This operationalmode enables the rocket motor to operate for the longest possiblemission time. As used herein, the term “steady state operation” isdefined as a state of the system after which the total thrust conditionshave been met and the command and actual thrust values are constant.

With all of the attitude control valves closed, higher operatingpressure within pressure vessel 14 and correspondingly higher thrust maybe accomplished by partially closing the axial thrust valve 10.Partially closing the axial thrust valve 10 will reduce the effectivecross-sectional area of the nozzle throat 48, resulting in a higheroperating pressure and therefore higher thrust. This will decrease themission time. As noted above, the axial thrust valve 10 may comprise apintle valve, with actuator 44, powered by battery 46 moving the pintleelement 50 toward and away from the nozzle throat 48 to change thenozzle throat area to alter pressure within the pressure vessel 14 andresulting thrust. While only a single axial thrust valve and associatedaxial thruster are depicted in the foregoing embodiment, it iscontemplated that more than one axial thrust valve and associated axialthruster may be employed without departing from the scope of the presentinvention. Maneuver control valves 28, 30, 36 a, 36 b, 38 a and 38 bmay, as with axial thrust valve 10, be actuated by battery-poweredactuators (not shown) powered by battery 46 or one or more otherbatteries. Alternatively, the valves, if electrically actuated, may bepowered by a fuel cell.

Thrust to any one of the maneuver control or axial thruster valves maybe controlled proportionately and substantially independently of thethrust provided to any other thruster valves. For example, the thrustprovided to maneuver control (pitch) thruster 32 by maneuver controlvalve 28 may be set to 100 lbf while all other maneuver control valvesare producing a negligible amount of thrust through their associatedthrusters. Then, to increase thrust in, for example, the yaw directionwhile maintaining the thrust in the pitch direction, maneuver control(yaw) valves 36 a and 38 a may be opened and maneuver control valves 30,36 b and 38 b may be closer further. By closing the maneuver controlvalves 30, 36 b and 38 b further, the pressure in pressure vessel 14 isincreased to increase mass flow. By opening maneuver control (yaw)valves 36 a and 38 a further, more mass flow is directed out of thosevalves into their associated yaw maneuver control thrusters 40 a and 42a, producing increased thrust.

Roll control may be achieved by opening two maneuver control valves,such as valves 36 a and 36 b to respectively power opposing, off-axismaneuver control thrusters 40 a and 40 b, produces offset thrust aboutlongitudinal axis L in a common plane transverse to the longitudinalaxis causing the vehicle to roll. With the addition of propellant massabove that which is required for axial thrust, maneuvering functions canthus be performed without affecting axial thrust levels. The maneuvercontrol thrusters may be smaller than the axial thruster 26, with eachmaneuver control thruster 32, 34 for pitch control and each maneuvercontrol thruster of the two sets of yaw and roll maneuver controlthrusters 40 a, 40 b and 42 a, 42 b providing a smaller force than theaxial thruster 26. For example, and not by way of limitation, maneuvercontrol thrusters 32 and 34 for pitch control may be designed to eachprovide 1,000 lbf maximum thrust capability, while maneuver controlthrusters 40 a, 40 b, 42 a and 42 b for yaw and roll control may each bedesigned to provide a 500 lbf maximum thrust capability.

Mass flow and, therefore, burn time, may also be controlledsubstantially independently of other system variables. For example, anull thrust and low mass flow scenario may be created by opening all ofthe valves to the point where all thrusts are offsetting and a minimumsteady state mass flow exists. To increase mass flow and keep maneuverthrust the same, all valves may be closed partially to increase pressurein the pressure vessel 14, thereby increasing mass flow. Minimization ofmass flow while meeting other system requirements is the generallypreferred operational state.

Changes in internal temperature will affect the pressure within thepressure vessel 14. Temperature as well as pressure sensors may be addedto the pressure vessel 14 to monitor these parameters, and the axialthrust valve flow area may be modulated to compensate for suchtemperature effects to achieve a substantially constant axial thrust, ifdesired. Flow through one or more maneuver control valves 28, 30, 36 a,36 b, 38 a and 38 b may also be modulated to affect pressure within thepressure vessel 14 to compensate for temperature effects, or to achievedesired thrust levels. The addition of pressure sensors (transducers) tothe pressure vessel 14 to monitor chamber pressure thereof is desirablesince factors other than temperature such as, for example, manufacturingvariations will affect system performance. The use of pressuretransducers enables modulation of the flow through the valvescommunicating with the pressure vessel 14 to compensate for any factorswhich affect chamber pressure. Feedback from the pressure transducersmay also be used in a closed loop control system to control desiredparameters of the propulsion system. Accelerometers may also be added tothe rocket motor to provide a more accurate measurement by which thrustmay be predicted or system performance monitored. Feedback from theaccelerometers may also be used in a closed loop control system tocontrol desired parameters of the propulsion system.

In another exemplary embodiment of the invention, additional maneuvercontrol valves may be used. Further, pitch and yaw maneuver controlthrusters may have mass flow provided thereto by the solid propellantgrain used to provide mass flow for the axial thruster and a separategas generator and associated thrusters may be provided for roll control.Alternatively, maneuver control thrusters for roll may be provided withmass flow by the solid propellant grain used to provide mass flow forthe axial thruster and a separate gas generator and associated thrustersprovided for pitch and yaw control. Finally, a pitch and roll or yaw androll thruster set may be provided with mass flow by the solid propellantgrain used for axial thrust, and the other maneuver control parameter,yaw or pitch, controlled by a separate system. However, due tofabrication and operational complexity as well as added vehicle weight,these alternatives are currently less preferred.

The maneuver control thrusters for pitch, yaw and roll may, instead ofbeing aimed transversely to the longitudinal axis L of the rocket motor,be oriented to release gases substantially in the direction of axialthrust (not shown). Thus, pitch, yaw and roll control thrusters may beindividually offset from the longitudinal axis L of the rocket motor;however, these maneuver control thrusters may, for example, be locatedand oriented to collectively form a concentric ring about thelongitudinal axis L of the rocket assembly, so that simultaneousoperation of certain or all of the associated maneuvering valves causesthe maneuver control thrusters to provide thrust to the vehicle withoutadjustment in pitch, yaw or roll. In such a configuration, and if themaneuver control thrusters may provide sufficient axial thrust, thepresence of a separate, main axial thrust valve to provide axial thrustis optional.

The thrust vector of the axial thruster 26 (see FIG. 1) may additionallybe altered by use of a gimbaled, rotationally movable nozzle exit coneto perform or assist in attitude control functions. As noted above, allmaneuver control and axial thrust valves may be proportional valves. Theproportional valves may be controlled electrically, pneumatically,hydraulically or mechanically and they may be linearly or nonlinearlyacting in their modes of operation. Each proportional valve may beconfigured to be shut down completely as well as to achieve a highthrust turn-down ratio.

FIG. 3 depicts, in schematic form, an exemplary configuration for arocket motor system including a rocket motor, in accordance with anembodiment of the present invention. A rocket motor system 200 includesa rocket motor controller 202 and a rocket motor 204, as describedherein. The rocket motor controller 202 operates on total thrust (ortotal propellant mass discharged from the valves per unit time), ratherthan on pressure only as in prior art control methodologies. The rocketmotor 204 may be configured as a solid propellant engine or as a hybridengine. The details of the structure of and suitable propellants,oxidizers and ignition sources are known to those of ordinary skill inthe art, and may also be found, for example, in U.S. Pat. No. 6,393,830,assigned to the assignee of the present invention and the disclosure ofwhich patent is incorporated herein by reference.

A solid rocket motor, according to an embodiment of the presentinvention, may comprise a pressure vessel 114 containing a suitablesolid propellant grain 122. Pressure vessel 114 is in selectivecommunication with an axial thruster 126 through axial thrust valve 110and in selective communication with a plurality of maneuver controlthrusters 130 for pitch, yaw and roll control through respectivelyassociated maneuver control valves 128. Any suitable number of maneuvercontrol valves 128 and associated maneuver control thrusters 130 may beemployed as desired or required, depending on the maneuver controlthruster layout chosen. Axial thrust valve 110 and maneuver controlvalves 128 may comprise proportional, or throttling type valves.

The axial thrust valve 110 and the maneuver control valves 128 areselectively controlled by rocket motor controller 202. In one embodimentof the present invention, the valves 110, 128 are configured asproportional valves responsive to proportional control by the rocketmotor controller 202. The rocket motor controller 202 computes the totalmass flow rate required to steer the rocket motor propelled vehicleincluding a rocket motor 204 configured as a solid propulsion system.The orientation of the rocket propelled vehicle including the rocketmotor system 200 is directly controlled by thrust generated byproportional valves 110, 128 coupled to the pressure vessel 114 inincluding propellent grain 122 and functioning as a gas generator. Oneimprovement over prior attempts includes minimization of the amount ofpropellant used to steer a rocket propelled vehicle while meeting netthrust demands. The rocket motor controller 202 calculates and generatessignals for controlling specific ones of valves 110, 128 to preferablycontrol undershoot and overshoot during command changes to theproportional valves 110, 128.

In order to provide control to the various valves, each of the valves isdesirably configured as a proportional valve that includes an orificewith a controllable flow area and is part of a valve system coupled to agas generator. The rocket motor controller 202 may be configured forcontrolling an even number of proportional valves 110, 128 forcontrolling the direction of thrust in both directions with acorresponding positively or negatively signed command or may be adaptedfor controlling an odd number of valves when a valve is not paired withan opposingly arranged complementary valve.

As used herein, the terms “propellant mass flow discharge rate,” “massflow discharged per unit time,” “time rate of change of dischargedpropellant mass,” and the like, may be abbreviated by the calculusnomenclature of “m-dot_(disch.)” Accordingly, such terms and phrases areused interchangeably throughout.

The control signals are calculated in rocket motor controller 202 bydefining a total thrust in terms of a time-rate of change of propellantmass, “m-dot_(disch)” discharged from all valves. While the rocket motorcontroller 202 is illustrated to include specific separate modules orsub-controllers, the separation is exemplary and not to be considered aslimiting of the scope of the various embodiments of the presentinvention. By way of example, the rocket motor controller 202 includes aconditioner controller 210, an automatic controller 212, anm-dot_(disch) controller 214, and a burn rate estimator or burn ratecontroller 216. Generally, the conditioner controller 210 calculates atotal thrust or total mass flow rate discharged command while them-dot_(disch) controller 214 establishes the total area required to meetthe total mass flow rate discharged command and distributes this area tothe valves in order to meet the net thrust commands.

Specifically, the rocket motor controller 202 includes a conditionercontroller 210 for calculating control signals based on total thrust ortotal propellant mass discharged from the valves per unit time asopposed to pressure only within the pressure vessel 114. In oneembodiment of the present invention as illustrated with respect to theflowchart of FIG. 4A, for each opposing valve pair, a net thrust commandis received 250 to calculate 252 a minimum total thrust (with aconfigurable, optional margin) required from the gas generator tosatisfy all net thrust commands. While not essential, calculation of aminimum total thrust ensures all net thrusts will be met, however,without calculation of the minimum total thrust, the actual total thrustmay be insufficient to meet net thrust requirements or more total thrustmay be commanded than required to meet net thrust requirements resultingin wasted propellant.

In FIG. 4A, the minimum total thrust is determined 254 herein as thesmallest thrust that will support all the net thrust requirements andcan be found as the maximum of either the sum of net thrusts 256required from each valve (after appropriate scaling based on valvesizes) or the worst case net thrusts 258 (after appropriate scalingbased on valve sizes). Both net thrust conditions may be consideredsince the total thrust requirements might not permit generatingsufficient net thrusts even though the total thrust is greater than thesum of the net thrust. This may occur because the specified total thrustdetermines the requisite pressure within the pressure vessel in order togenerate the total thrust. This resulting pressure may be insufficientto generate the required net thrust, (i.e., even with one valve wideopen and its complementary valve closed, insufficient net thrust may begenerated). Thus, the total thrust command may be overridden if morethrust is needed to satisfy net thrust requirements.

In another embodiment of the present invention as illustrated withrespect to FIG. 4B, total thrust command and net thrust commands can beindependent inputs 206 (FIG. 3) to the conditioner controller 210 ratherthan using net thrust commands to calculate the total thrust command. Inthe present embodiment, no attempt to override the total thrust commandis necessary. Furthermore, the total thrust command is received 260 andthe net thrust commands are also received 262. Since the total thrustcommand inputs 206 are input independently, the net thrust commandsshould not be higher than the total thrust commanded.

The total thrust command is determined 264 as the sum of the net thrusts266 or an approximation of the sum of the net thrusts less a difference268 corresponding to a deficient amount. Accordingly, when deployed, thenet thrusts should be achieved or at least approximated as closely aspossible. In a practical application, the net thrust may not be metbecause the pressure associated with the total thrust command isinsufficient to meet the net thrust commands. An override conditionwithin the controller may occur during transition (increases anddecreases) of total thrust beyond a prescribed value and during pressureemergencies and startup conditions. The override may be in control for ashort predetermined duration after the pressure vessel's pressureapproximately reaches the pressure necessary to generate the totalthrust.

In either of the embodiments, the total thrust command should be met aslong as the total thrust is not beyond the pressure limits of thedeployed system, as illustrated with respect to the flowchart of FIG. 5.In FIG. 5, the gas generator is commanded 270 according to the totalthrust command. If a measured 272 pressure in the gas generator exceedsa configurable value (e.g., high pressure limit), all valves will becommanded fully open 274, disregarding all thrust commands, until thepressure drops below the medium pressure limit. If a measured 272pressure in the gas generator does not exceed a configurable value(e.g., high pressure limit), the pressure is further compared. If ameasured 276 pressure in the gas generator exceeds a configurable value(e.g., medium pressure limit), the respective valves are incrementallyopened 278.

FIG. 6 is a flowchart illustrating valve command mode options inresponse to thrust commands from the conditioner controller 210 (FIG.3), in accordance with various embodiments of the present invention.According to a specific mode defined herein by a mode configuration 280,the axial thrust valve or divert valves 110 may completely close,according to a minimal response time mode 282, until the total thrustcommand can be achieved to minimize total thrust response time byproviding for the most rapid change in pressure. Alternatively, theaxial thrust valve or divert valves 110 may react according to asustained-gradual pressure changing mode 284, causing a lessinstantaneous rate of change resulting in a less than minimized responsetime due to a slower rate of pressure change.

In one embodiment, the response 286 is based on the slope of the changein the total thrust command. If there is a sudden change in the totalthrust command so that the command changes for one time interval toanother by an amount determined by a predefined constant, then the axialthrust valve or divert valves 110 may be completely closed or opened 288until the total thrust command is achieved (i.e., minimal total thrustresponse time by providing for the most rapid change in pressure). Ifthe changes are slower, as would be the case if the thrust command wasramped, then control will be accomplished by an automatic controller 212(i.e., the maneuver control valves 128 react in a continuous manner).During changes in total thrust commands, the maneuver control valves 128try to meet the net thrust commands since the maneuver control valves128 are not commanded fully open or fully closed to meet the totalthrust command when a non-zero axial thrust valve or divert valves 110command is present.

During transient operation when the change in total thrust command isgreater than a defined threshold value, the axial thrust valve or divertvalves 110 is completely closed 288 if the total thrust command hasincreased. Conversely, if the total thrust command has decreased, theaxial thrust valve or divert valves 110 is completely opened 288. Alsoduring the transient operation with the axial thrust valve or divertvalves 110 either completely closed or completely open, the maneuvercontrol valves 128 may or may not remain in either an open or closedposition until the pressure is attained 290 within the pressure vessel114 which corresponds to a percentage of the newly commanded totalthrust. When this pressure is reached, the valves 110, 128 move to aposition, for a brief configurable duration to allow the valvepositioning to settle 292, corresponding to a calculated area requiredto meet the total thrust command.

As illustrated with respect to FIG. 7, after a configurable duration, anautomatic controller 212 (FIG. 3) within rocket motor controller 202 isengaged after achieving the necessary pressure for total thrust andafter a configurable time has elapsed for the valves to settle, asillustrated above with respect to FIGS. 4-6. Generally, the automaticcontroller 212 modifies the total area so that the computedm-dot_(disch) matches the m-dot corresponding to the desired totalthrust. It also is used to provide m-dot_(disch), control whenever thechange in the commanded m-dot is small and for rejection of anydisturbance that might cause m-dot_(disch) to change.

When a change in total thrust command occurs and is smaller than thethreshold 294 value referenced above, the automatic controller 212remains active (i.e., no conditioner controller 210 override 296occurs). The automatic controller 212 is configured to controlm-dot_(disch); however, the present embodiments do not require a sensorfor measuring the m-dot_(disch). Instead, the value of m-dot_(disch)that is used for feedback control is based on a first-principlecomputation. The total thrust command is converted 298 to anm-dot_(disch) command which is then compared with the computedm-dot_(disch). The difference is used to drive the automatic controller212 to produce a new total valve area so that the computed m-dot_(disch)agrees with the commanded m-dot_(disch).

Continuing, the automatic controller 212 converts 298 the total thrustcommand into an equivalent m-dot_(disch) command. The logarithm of them-dot_(disch) command is computed 300 and the logarithm of the computedm-dot_(disch) is also computed 302. An error is generated 304 from thesequantities which is fed into an m-dot_(disch) controller 214 (FIG. 3).The logarithm of total area for all valves is computed 306 andexponentiated 308 to compute the total area required by all valves tomeet the required m-dot_(disch) and therefore the total thrust command.

The relationship of the valve area to m-dot_(disch) is highly nonlinearin pressure and free volume and therefore a controller that isconfigured for computation of logarithms of m-dot_(disch) and totalvalve area is desirable. Use of logarithms linearizes the overallsteady-state gain between area and m-dot_(disch). By way ofimplementation of one or more embodiments, linearization permits the useof a motor controller capable of working over a wide range of pressuresand free volumes. Without such linearization, the controller would needto be much more complex, possibly requiring switching between multiplecontrollers.

If a conflict exists between meeting total thrust and net thrusts, apriority 310 is assigned for both the axial thruster valve 110 andmaneuver control valves 128. In one embodiment, the priority is assignedsuch that maneuver control valves 128 give highest priority to meetingnet thrust commands and the axial thrust valve 110 gives highestpriority to meeting total thrust. Once the total valve area is computed306, the total valve area is algebraically distributed 312 among thevarious valves to meet net thrusts and total thrust requirements. Thecontroller attempts to achieve the commanded total thrust, butacquiesce, if need be, to maintain as much of the maneuver controlthrust valves 128 net thrusts commands as possible.

As illustrated with respect to FIG. 8, a burn rate estimator or burnrate controller 216 (FIG. 3) dynamically calculates 314 the actual burnrate based on total valve area and gas generator pressure. Knowledge ofthe actual burn rate is beneficial for operating the system efficientlyduring changes in total thrust commands and predicting propellantburnout, such as in the case of mission planning. Additionally, anaccurate burn rate controller provides improved m-dot_(disch)predictions and as a result, improved control. The burn rate estimate ispreferably performed when the system is in a steady state. During thesteady state operation, the burn rate may be adjusted 318 so thepredicted estimate of the pressure vessel pressure matches 316 themeasured pressure vessel pressure.

According to the various embodiments of the present invention, it isdesirable to control the product of the valve area and the pressurerather than just the pressure alone since such a product determines thethrust and is proportional to m-dot_(disch). Accordingly, in a pressureonly approach, if the burn rate is higher than expected, the throat areawill be higher than expected to control pressure resulting in generationof more total m-dot than required which further results in wastedpropellant. If the burn rate is lower than expected, controllingpressure does not guarantee that the thrust command will be achieved.

Accordingly, through utilization of m-dot_(disch) control as describedherein with regard to the various embodiments of the present invention,when the burn rate is higher than expected, the valve area will behigher, but the pressure will be lower to meet the m-dot command, thusconserving propellant. If the burn rate is lower than expected, thepressure will be increased to meet the thrust commands.

While the present invention has been disclosed in terms of certainexemplary embodiments, those of ordinary skill in the art will recognizeand appreciate that the invention is not so limited. Additions,deletions, and modifications to the disclosed embodiments may beeffected without departing from the scope of the invention as claimedherein. Similarly, features from one embodiment may be combined withthose of another while remaining within the scope of the invention.

1. A method of controlling thrust in a rocket motor, comprising: positioning valves of an energized rocket motor to achieve an initial total valve area corresponding to a total thrust command; determining a commanded discharged propellant mass flow discharge rate that corresponds to the total thrust command; calculating a compensated total valve area corresponding to an error between the commanded propellant mass flow discharge rate and at least one of a calculated propellant mass flow discharge rate and a measured propellant mass flow discharge rate; and repositioning at least one of the valves to achieve the compensated total valve area.
 2. The method of claim 1, further comprising calculating a the total thrust command from at least one of a sum of all net thrust commands of each of the valves of the energized rocket motor and a maximum net thrust command.
 3. The method of claim 1, further comprising: measuring a pressure within a gas generator of the rocket motor; and repositioning the valves according to an override command when the pressure exceeds a pressure limit of the gas generator.
 4. The method of claim 1, wherein repositioning at least one of the valves further comprises positioning at least one of the valves to one of a substantially maximally open and a substantially maximally closed position in response to the error between the commanded propellant mass flow discharge rate and at least one of the calculated propellant mass flow discharge rate and the measured propellant mass flow discharge rate.
 5. The method of claim 4, further comprising repositioning at least one of the valves incrementally to approach the initial total valve area when a measured pressure within a gas generator of the rocket is within a threshold corresponding to the total thrust command.
 6. The method of claim 1, wherein repositioning at least one of the valves further comprises incrementally moving at least one of the valves.
 7. The method of claim 1, further comprising repositioning at least one of the valves according to a priority command when the total thrust command exceeds available thrust of the rocket motor.
 8. The method of claim 1, further comprising: calculating a burn rate from a measured pressure within a gas generator of the rocket motor and the compensated total valve area; and estimating the burn rate utilizing a comparison between the measured pressure and an estimated pressure.
 9. A method for controlling thrust in a rocket motor, comprising: calculating a minimum total thrust required from a gas generator of a rocket motor to meet at least one net thrust command; positioning axial and maneuver control valves of the rocket motor according to the minimum total thrust; and repositioning at least one of the axial and maneuver control valves to change a valve area of the axial and maneuver control valves in response to a comparison of a commanded propellant mass flow discharge rate and a calculated propellant mass flow discharge rate.
 10. The method of claim 9, wherein calculating the minimum total thrust comprises calculating at least is-one of a sum of the net thrust commands and a sum of worst case net thrusts.
 11. The method of claim 9, wherein repositioning at least one of the axial and maneuver control valves comprises positioning at least one of the axial and maneuver control valves to one of a substantially maximally open and a substantially maximally closed position in response to the minimum total thrust command.
 12. The method of claim 9, wherein repositioning at least one of the axial and maneuver control valves comprises incrementally moving at least one of the axial and maneuver control valves in response to the minimum total thrust command.
 13. The method of claim 9, further comprising calculating the commanded propellant mass flow discharge rate from the total thrust command.
 14. The method of claim 9, further comprising repositioning at least one of the axial and maneuver control valves according to a priority command when the total thrust command exceeds available thrust of the rocket motor.
 15. The method of claim 9, further comprising: calculating a burn rate from a measured pressure within a gas generator of the rocket motor and a total valve area; and recalculating the burn rate utilizing a comparison between the measured pressure and an estimated pressure. 